ULA IVF and their H2-O2 Thruster System program implications for deep space Cubesat missions

December 18, 2014

ULA IVF program was meant to replace the Helium ullage gas system and provide station keeping using residual fuel,does the H2-O2 Thruster System program replace the IVF program? NASA has partly defunded fuel depot work and transferred this work to a ground test program only

The author of paper, “orbital disposal of launch vehicle upper stages” linkeDin page shows her not in the project lead as of August 2014.Should we be sad? Yes, but maybe not to worry ,the paper does state that technology useful in ACES might make its way to the SLS upper stage(with 4 engines) and upper stages with fewer than 4 engines.My take on this is that the air force with its recent RFI for engines and NASA contracted SLS upper stage with Boeing might yet create synergies for both programs.

A number of years ago I mused with Mr Bernard Kutter of ULA about the idea of keeping LHe 2 cool with a fuel depot.The LHE 2 was for transfer to cryogenic space telescope. Later I discarded the idea in favor of a fuel tanker delivering the cryogenic helium to space telescope and then making its fuel depot run.

On selenium boondocks I posted comments on ideas of hybrid chemical ion hybrid in space stages but now I would like to write more about disposal burns that could be further optimized with a cryogenic Xenon in space powered ion engine,And Mr Goff hs written about early SLS upper stages in regards to ACES.

If Ms Gravelee is no longer ULA chief of the fuel depot program there I hope she is on the SLS ACES/EUS work.

The Boeing SLS would be the way to create a SLS upper stage with the ability that has been advocated for the ACES stage

We propose that IVF and/or H2-O2 thruster program be augmented with a LXe tank kept cryogenic with LH2 boil off. An ion engine could loft all upper stages into Heliocentric orbits and turn all upper stages into cubesat /secondary payload opportunities

99-0288 This paper describes multiple burns over many months from GTO and escape

1. “At some predetermined time after
GTO launch, perform the first burn at GTO
perigee in the orbit plane to enter a high ellipse
of between 700,000 km to 1,500,000 km
apogee distance. This will require 720 to 750
m/s delta-v out of the GTO” (3)

2.”To deorbit a 2727 kg spent stage inserted into the GSO belt would require some 1300 kg of
propellant, seemingly not an unreasonable mass”(1)

A question we have is how much fuel or m/s delta-v is required to send 2727 Kg into a Moon crossing orbit? another would be how much LXe would be needed for the same mission?It is 1,500 M/S from GTO to earth so according to Gravlee 1,300 kg of propellant is required to provide 1,500 m/s to a 2,727 kg stage to destructive reentry. but we want to preserve spent stages for future use so heliocentric or L2 storage is better.We need 760 M/S to L-1 Halo so this must be half of the 1,300 kg**.GTO to translunar is 700 M/S

**** “Raise the apogee to about 1 million km, so
hat the apogee velocity is very small. This
adds only about 72 m/s to the ∆v compared
with an apogee raise to the Moon’s distance.
– Perform a plane change at apogee (∆v =
300 m/s approximately), so that the orbit’s
return leg meets the Moon’s orbit”